1. Field of the Invention
This invention pertains to a heat transfer device and in particular to a device for transferring dissipated heat from a pressure vessel in a spacecraft.
2. Description of the Related Art
The economical transfer of dissipated heat from pressure vessels within spacecraft is an existing problem. One such heat source within a spacecraft are the electrochemical batteries which supply spacecraft power during eclipse periods when the solar arrays are not illuminated. As the electrochemical batteries are discharged and overcharged, the battery dissipates energy in the form of heat. Particularly, thermal control to maintain the temperature of such heat generating devices as the electrochemical battery within a desirable operating temperature (typically 0 to 20 degrees Celsius (.degree.C.)) is accomplished for small spacecraft electrical loads through heat transfer via a battery mounting structure. However, for large thermal loads, additional thermal control devices are required to remove the excess heat generated.
Generally, in the space environment, thermal control of a spacecraft battery is achieved by conduction from the battery pressure vessel to a radiator, then by direct or indirect radiation to space. All of these methods require a means for securing a cylindrical pressure vessel to a flat mounting surface. The most common method is to secure the pressure vessel to a radiator panel by the use of an aluminum collar to conduct the heat from the pressure vessel wall through the aluminum collar to a radiator panel. Under large dissipative heat loads, the aluminum collar must be very large in order to conduct the high heat load and to maintain a uniform temperature across the pressure vessel. Heat pipes and louvers may be added to the thermal design to reduce the mass of the aluminum collar, but overall weight, complexity, and cost of the design is increased.
A previous design consisted of a mounting system external to the spacecraft whereby aluminum brackets linked a magnesium battery heat sink to a close-out panel on the spacecraft. The battery is restrained by the heat sink in the lateral axes while stainless steel straps restrain the axial displacement. This type of design was not faced with a weight limitation and as a result of the heat conduction requirements was quite heavy. In this design the magnesium heat sink was required to be isolated from the aluminum brackets, therefore, heat can not be transferred from the heat sink through the brackets into the spacecraft. The heat dissipated from the battery was radiated into space by radiator panels. Use of this system is impractical when interior mounting of the battery is required because the heat would be rejected to the inside of the spacecraft thereby requiring a heavier and more expensive thermal control system.